EP3705704A1 - Wellenvorrichtung für ein gasturbinentriebwerk - Google Patents

Wellenvorrichtung für ein gasturbinentriebwerk Download PDF

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Publication number
EP3705704A1
EP3705704A1 EP20157725.1A EP20157725A EP3705704A1 EP 3705704 A1 EP3705704 A1 EP 3705704A1 EP 20157725 A EP20157725 A EP 20157725A EP 3705704 A1 EP3705704 A1 EP 3705704A1
Authority
EP
European Patent Office
Prior art keywords
shaft
fan
shaft portion
ratchet
gas turbine
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
EP20157725.1A
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English (en)
French (fr)
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EP3705704B1 (de
Inventor
Michael Booth
Steven Culwick
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Rolls Royce PLC
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Rolls Royce PLC
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Publication date
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Publication of EP3705704A1 publication Critical patent/EP3705704A1/de
Application granted granted Critical
Publication of EP3705704B1 publication Critical patent/EP3705704B1/de
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16D—COUPLINGS FOR TRANSMITTING ROTATION; CLUTCHES; BRAKES
    • F16D43/00—Automatic clutches
    • F16D43/02—Automatic clutches actuated entirely mechanically
    • F16D43/20—Automatic clutches actuated entirely mechanically controlled by torque, e.g. overload-release clutches, slip-clutches with means by which torque varies the clutching pressure
    • F16D43/202—Automatic clutches actuated entirely mechanically controlled by torque, e.g. overload-release clutches, slip-clutches with means by which torque varies the clutching pressure of the ratchet type
    • F16D43/2022—Automatic clutches actuated entirely mechanically controlled by torque, e.g. overload-release clutches, slip-clutches with means by which torque varies the clutching pressure of the ratchet type with at least one part moving axially between engagement and disengagement
    • F16D43/2024—Automatic clutches actuated entirely mechanically controlled by torque, e.g. overload-release clutches, slip-clutches with means by which torque varies the clutching pressure of the ratchet type with at least one part moving axially between engagement and disengagement the axially moving part being coaxial with the rotation, e.g. a gear with face teeth
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/36—Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/40—Movement of components
    • F05D2250/41—Movement of components with one degree of freedom
    • F05D2250/411—Movement of components with one degree of freedom in rotation
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/40—Transmission of power
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2270/00—Control
    • F05D2270/01—Purpose of the control system
    • F05D2270/09—Purpose of the control system to cope with emergencies
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present disclosure relates to shaft apparatus for gas turbine engines, and more particularly to seizure protection for fan shafts in gas turbine engines.
  • a shaft apparatus for a gas turbine engine comprising a first shaft portion, a second shaft portion, and a ratchet mechanism configured to permit the first shaft portion to rotate with respect to the second shaft portion in a first direction, and to prevent the first shaft portion from rotating with respect to the second shaft portion in a second direction opposite to the first direction
  • the ratchet mechanism comprises a first ratchet element configured to rotate with the first shaft portion, and a second ratchet element configured to rotate with the second shaft portion
  • each of the first and second ratchet elements comprise a respective ratchet surface for engagement with the ratchet surface of the other of the first or second ratchet element; and, the ratchet mechanism is configured such that downstream axial movement of the first shaft portion relative to the second shaft portion causes the ratchet surfaces of the first and second ratchet elements to be disengaged.
  • the ratchet mechanism may be configured such that: a) the first shaft portion is permitted to rotate faster than the second shaft portion in the first direction; b) the second shaft portion may not rotate faster than the first shaft portion in the first direction (i.e. the first shaft portion must rotate at least as fast as the second shaft portion in the first direction); and c) the first shaft portion may not rotate relative to the second shaft portion in the second direction (i.e. if the first shaft portion is rotating in the second direction, then the second shaft portion must also be rotating at the same speed in the second direction).
  • the ratchet mechanism may permit the second shaft portion to apply a torque to the first shaft portion, but prevent the first shaft portion from applying a torque to the second shaft portion.
  • the shaft apparatus may be a fan shaft apparatus.
  • the first shaft portion may be configured for attachment to a fan of a gas turbine engine.
  • the second shaft portion may be configured to apply a driving torque to the first shaft portion in the first direction so as to drive a fan attached to the first shaft portion.
  • the ratchet mechanism may comprise a first ratchet element configured to rotate with the first shaft portion, and a second ratchet element configured to rotate with the second shaft portion.
  • the first ratchet element may be arranged axially downstream of the second ratchet element.
  • the axial direction may refer to any direction parallel to a rotational axis of the first and/or second shaft portions, or may refer to any direction parallel to the principal rotational axis of the gas turbine engine. Downstream in the axial direction generally refers to the direction of the exhaust end of the engine, while upstream refers generally to the direction of the intake end of the engine.
  • the ratchet surface of the first ratchet element may be configured to face at least partially upstream in an axial direction, and wherein the ratchet surface of the second ratchet element is configured to face at least partially downstream in an axial direction.
  • the ratchet mechanism may comprise a curvic coupling.
  • the ratchet mechanism may comprise a clutch.
  • a gas turbine engine for an aircraft comprising a shaft apparatus according to the first aspect described above.
  • the gas turbine engine may further comprise an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades and being attached to the first shaft portion; and a gearbox that receives an input from the core shaft and outputs drive to the fan via the second shaft portion so as to drive the fan at a lower rotational speed than the core shaft.
  • the turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft.
  • the engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor.
  • the second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.
  • a method of providing seizure protection in an engine comprising a fan and a torque shaft for driving the fan in a first direction comprising permitting the fan to rotate relative to the torque shaft in the first direction, and preventing the fan from rotating relative to the torque shaft in a second direction opposite to the first direction.
  • a shaft apparatus may be provided to permit and prevent relative rotation of the fan with respect to the torque shaft in the first and second directions respectively.
  • the torque shaft may comprise or may be the second shaft portion.
  • the fan may comprise or may be attached to the first shaft portion.
  • Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor.
  • a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core.
  • the gas turbine engine may comprise a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.
  • the input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear.
  • the core shaft may rigidly connect the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).
  • the gas turbine engine as described and/or claimed herein may have any suitable general architecture.
  • the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts.
  • the turbine connected to the core shaft may be a first turbine
  • the compressor connected to the core shaft may be a first compressor
  • the core shaft may be a first core shaft.
  • the engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor.
  • the second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.
  • the second compressor may be positioned axially downstream of the first compressor.
  • the second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor.
  • the gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above).
  • the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above).
  • the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above.
  • the gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used.
  • the gearbox may be a "planetary” or “star” gearbox, as described in more detail elsewhere herein.
  • the gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of from 3 to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2.
  • the gear ratio may be, for example, between any two of the values in the previous sentence.
  • the gearbox may be a "star” gearbox having a ratio in the range of from 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges.
  • a combustor may be provided axially downstream of the fan and compressor(s).
  • the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided.
  • the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided.
  • the combustor may be provided upstream of the turbine(s).
  • each compressor may comprise any number of stages, for example multiple stages.
  • Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence may be variable).
  • the row of rotor blades and the row of stator vanes may be axially offset from each other.
  • each turbine may comprise any number of stages, for example multiple stages.
  • Each stage may comprise a row of rotor blades and a row of stator vanes.
  • the row of rotor blades and the row of stator vanes may be axially offset from each other.
  • Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or 0% span position, to a tip at a 100% span position.
  • the ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25.
  • the ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e.
  • the values may form upper or lower bounds), for example in the range of from 0.28 to 0.32. These ratios may commonly be referred to as the hub-to-tip ratio.
  • the radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade.
  • the hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.
  • the radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge.
  • the fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165 inches).
  • the fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 240 cm to 280 cm
  • the rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm.
  • the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 330 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm.
  • the fan In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity U tip .
  • the work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow.
  • a fan tip loading may be defined as dH/U tip 2 , where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and U tip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed).
  • the fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all values being dimensionless).
  • the fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.
  • Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions.
  • the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20.
  • the bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of form 12 to 16, 13 to 15, or 13 to 14.
  • the bypass duct may be substantially annular.
  • the bypass duct may be radially outside the core engine.
  • the radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.
  • the overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor).
  • the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75.
  • the overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 50 to 70.
  • Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg -1 s, 105 Nkg -1 s, 100 Nkg -1 s, 95 Nkg -1 s, 90 Nkg -1 s, 85 Nkg -1 s or 80 Nkg -1 s.
  • the specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 80 Nkg -1 s to 100 Nkg -1 s, or 85 Nkg -1 s to 95 Nkg -1 s.
  • Such engines may be particularly efficient in comparison with conventional gas turbine engines.
  • a gas turbine engine as described and/or claimed herein may have any desired maximum thrust.
  • a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN.
  • the maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).
  • a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 330kN to 420 kN, for example 350kN to 400kN.
  • the thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3kPa, temperature 30 degrees C), with the engine static.
  • the temperature of the flow at the entry to the high pressure turbine may be particularly high.
  • This temperature which may be referred to as TET
  • TET may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane.
  • the TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K.
  • the TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).
  • the maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K.
  • the maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 1800K to 1950K.
  • the maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.
  • MTO maximum take-off
  • a fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials.
  • at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre.
  • at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material.
  • the fan blade may comprise at least two regions manufactured using different materials.
  • the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade.
  • a leading edge may, for example, be manufactured using titanium or a titanium-based alloy.
  • the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.
  • a fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction.
  • the fan blades may be attached to the central portion in any desired manner.
  • each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc).
  • a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc.
  • the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring.
  • any suitable method may be used to manufacture such a bladed disc or bladed ring.
  • at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.
  • variable area nozzle may allow the exit area of the bypass duct to be varied in use.
  • the general principles of the present disclosure may apply to engines with or without a VAN.
  • the fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades.
  • cruise conditions have the conventional meaning and would be readily understood by the skilled person.
  • the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise of a given mission (which may be referred to in the industry as the "economic mission") of an aircraft to which the gas turbine engine is designed to be attached.
  • mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint - in terms of time and/or distance - between top of climb and start of descent.
  • Cruise conditions thus define an operating point of, the gas turbine engine that provides a thrust that would ensure steady state operation (i.e.
  • cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide - in combination with any other engines on the aircraft - steady state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach Number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude).
  • the mid-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined.
  • the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.
  • the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 10000m to 15000m, for example in the range of from 10000m to 12000m, for example in the range of from 10400m to 11600m (around 38000 ft), for example in the range of from 10500m to 11500m, for example in the range of from 10600m to 11400m, for example in the range of from 10700m (around 35000 ft) to 11300m, for example in the range of from 10800m to 11200m, for example in the range of from 10900m to 11100m, for example on the order of 11000m.
  • the cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.
  • the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30kN to 35kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (11582m).
  • the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 50kN to 65kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000ft (10668m).
  • a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein.
  • cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust.
  • an aircraft comprising a gas turbine engine as described and/or claimed herein.
  • the aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.
  • a method of operating a gas turbine engine as described and/or claimed herein may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number).
  • a method of operating an aircraft comprising a gas turbine engine as described and/or claimed herein.
  • the operation according to this aspect may include (or may be) operation at the mid-cruise of the aircraft, as defined elsewhere herein.
  • FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9.
  • the engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B.
  • the gas turbine engine 10 comprises a core 11 that receives the core airflow A.
  • the engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20.
  • a nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18.
  • the bypass airflow B flows through the bypass duct 22.
  • the fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.
  • the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place.
  • the compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted.
  • the resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust.
  • the high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27.
  • the fan 23 generally provides the majority of the propulsive thrust.
  • the epicyclic gearbox 30 is a reduction gearbox.
  • FIG. 2 An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Figure 2 .
  • the low pressure turbine 19 (see Figure 1 ) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30.
  • a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30 Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34.
  • the planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis.
  • the planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9.
  • an annulus or ring gear 38 Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.
  • low pressure turbine and “low pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23).
  • the "low pressure turbine” and “low pressure compressor” referred to herein may alternatively be known as the "intermediate pressure turbine” and “intermediate pressure compressor”. Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.
  • the epicyclic gearbox 30 is shown by way of example in greater detail in Figure 3 .
  • Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in Figure 3 .
  • Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.
  • the epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed.
  • the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38.
  • the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.
  • any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10.
  • the connections (such as the linkages 36, 40 in the Figure 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility.
  • any suitable arrangement of the bearings between rotating and stationary parts of the engine may be used, and the disclosure is not limited to the exemplary arrangement of Figure 2 .
  • the gearbox 30 has a star arrangement (described above)
  • the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in Figure 2 .
  • the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.
  • gearbox styles for example star or planetary
  • support structures for example star or planetary
  • input and output shaft arrangement for example star or planetary
  • bearing locations for example star or planetary
  • the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).
  • additional and/or alternative components e.g. the intermediate pressure compressor and/or a booster compressor.
  • gas turbine engines to which the present disclosure may be applied may have alternative configurations.
  • such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts.
  • the gas turbine engine shown in Figure 1 has a split flow nozzle 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core engine nozzle 20.
  • this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle.
  • One or both nozzles may have a fixed or variable area.
  • the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example.
  • the gas turbine engine 10 may not comprise a gearbox 30.
  • the geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in Figure 1 ), and a circumferential direction (perpendicular to the page in the Figure 1 view).
  • the axial, radial and circumferential directions are mutually perpendicular.
  • Figures 4a and 4b schematically show a generic fan 400 and torque shaft 402 of a gas turbine engine in normal, driven operation.
  • Figure 4a shows a side view of the system to illustrate the axial forces
  • Figure 4b shows a front view of the system along the axis of the torque shaft to illustrate the rotational aspects of the system, such as torques.
  • the turbines of the engine apply a driving torque to the torque shaft 402 (either directly or via a gearbox as described above), and the fan 400 is attached to the torque shaft 402 such that the fan is driven in the direction of arrow F in Figure 4b .
  • Arrow D in Figure 4b represents the driving torque applied by the torque shaft 402 to the fan 400.
  • a reaction torque (as represented by arrow R) will be applied to the fan 400 and the shaft 402 by the aerodynamic resistance of the fan 400 as it is rotated.
  • the torque shaft 402 applies a driving torque to the fan in a first rotational direction, represented by the arrow F.
  • the fan 400 when driven, generates a forward axial thrust T represented by arrows, which is applied to the torque shaft 402.
  • This thrust force is reacted by a thrust bearing 404 as a reaction force r.
  • the torque shaft 402 experiences a tensile axial load, as represented by arrow t.
  • Figures 5a and 5b schematically show the same generic fan 400 and torque shaft 402 in windmilling operation.
  • Figure 5a shows a side view of the system to illustrate the axial forces
  • Figure 5b shows a front view of the system along the axis of the torque shaft to illustrate the rotational aspects of the system, such as torques.
  • the fan 400 when windmilling, is pushed axially rearward by the oncoming air, which generates a rearward axial force T' represented by arrows, which is applied to the torque shaft 402.
  • This force is reacted by a thrust bearing 404 as a reaction force r'.
  • the torque shaft 402 experiences a compressive axial load, as represented by arrow C.
  • Figure 6 an exemplary shaft apparatus 100 is shown in a gas turbine engine 10.
  • Figure 6 shows a schematic cross sectional view of at upper half of the shaft apparatus 100 about the main rotational axis A of the shaft apparatus and, more generally, the gas turbine engine.
  • the shaft apparatus 100 comprises a first shaft portion 102 and a second shaft portion 104.
  • the first and second shaft portions 102,104 are connected by a ratchet mechanism 106.
  • the ratchet mechanism 106 is configured to permit the first shaft portion 102 to rotate with respect to the second shaft portion 104 in a first direction, and to prevent the first shaft portion 102 from rotating with respect to the second shaft portion 104 in a second direction opposite to the first direction, as will be described in more detail below.
  • the shaft apparatus 100 is a fan shaft apparatus.
  • the first shaft portion 102 is attached to a fan 108.
  • the first shaft portion is configured to rotate about the axis A such that the fan 108 can be rotated therewith.
  • the first shaft portion 102 may be secured in the axial direction by a thrust bearing (not shown).
  • the second shaft portion 104 is configured to apply a driving torque to the first shaft portion 102, and may be referred to as a "torque shaft".
  • the second shaft portion 104 is configured to apply a driving torque to the first shaft portion 102 in a first direction F' (as illustrated by the arrow and 'out of page' dot on the fan 108) about the axis A so as to drive a fan attached to the first shaft portion 102 in the direction F'.
  • the direction F' is clockwise when viewed from the front along the axis A (i.e. the same as direction F in Figs 4b and 5b ).
  • the second shaft portion 104 in this example is driven by the output of a power gear box 110, for example a gear box as described in relation to figures 1, 2 , and 3 above. It should be understood that the principles of the shaft apparatus could also be applied to gas turbine engines without gear boxes, where the second shaft portion could be driven directly by a turbine.
  • the second shaft portion comprises a thrust bearing 112 which isolates the gear box 110 from axial loads on the second shaft portion.
  • the ratchet mechanism 106 is configured such that the second shaft portion 104 may not rotate faster than the first shaft portion 102 in the direction F. Accordingly, as the rotation speed of the second shaft portion 104 in direction F' increases, so must the rotation speed of the first shaft portion 102 and the fan 108 (i.e. the first shaft portion 102 must rotate at least as fast as the speed of the second shaft portion 104 in the first direction). This is necessary for the second shaft portion 104 to be able to drive the first shaft portion 102.
  • the ratchet mechanism 106 does however permit the first shaft portion 102 to rotate faster than the second shaft portion 104 in the first direction. For example, if the second shaft portion is static, rotating in the opposite direction to F', or rotating in the direction F' more slowly than the first shaft portion 102, then the first shaft portion 102 and the fan 108 can rotate relative to the second shaft portion 104 in direction F'.
  • the ratchet mechanism 106 prevents the first shaft portion 102 from rotating relative to the second shaft portion 104 in a direction opposite to F' (i.e. if the first shaft portion is rotating in the opposite direction to F', then the second shaft portion must also be rotating at the same speed in the that direction).
  • the fan 108 and thus the first shaft portion 102 would not typically need to rotate in the opposite direction to F' in driven or windmill conditions.
  • the shaft apparatus 100 is therefore configured such that in normal driven use (per Figure 4 above) the first shaft portion 102 and therefore the fan 108 can be driven normally by the second shaft portion 104.
  • the first shaft portion 102 and thus the fan 108 is permitted by the ratchet mechanism 106 to windmill in the direction F' relative to the second shaft portion 104.
  • the drag of the engine can be minimised in an engine failure scenario by virtue of the shaft mechanism 100, maintaining better aircraft control.
  • the ratchet mechanism 106 comprises a first ratchet element 114 configured to rotate with the first shaft portion 102, and a second ratchet element 116 configured to rotate with the second shaft portion 104.
  • Each of the ratchet elements 114,116 comprise a respective ratchet surface 118 for engagement with the ratchet surface 118 of the other ratchet element 114,116.
  • the ratchet mechanism 106 could be any apparatus or assembly which permits relative rotation of the first and second shaft elements in one direction but not the opposing direction.
  • the ratchet elements 114,116 could be curvic couplings (see Figures 8 and 9 below) or clutch elements, such as directional clutch elements.
  • FIG. 7a and 7b an alternative engine 20 having a further exemplary shaft apparatus 200 is shown. Comparable features between the shaft apparatus 100 and the shaft apparatus 200 are indicated by reference numerals differing by 100.
  • the first ratchet element 214 of the first shaft element 202 is arranged axially downstream of the second ratchet element 216 of the second shaft element 204, for reasons which shall be explained further below.
  • the ratchet surface 218' of the first ratchet element 214 is configured to face at least partially upstream in the axial direction and the ratchet surface 218" of the second ratchet element 216 is configured to face at least partially downstream in the axial direction.
  • the ratchet surfaces 218',218" are both perpendicular to the axis A, but it should be understood that in other examples, the ratchet surfaces could be arranged obliquely to the axis A and achieve similar effects.
  • the arrangement of the ratchet mechanism 206 of the shaft apparatus 200 is particularly advantageous when considering the forces at play in the system in the driven and windmill conditions as described in relation to Figures 4 and 5 above.
  • the fan 208 urges the first shaft portion 202 axially forward, meaning that the forward-facing ratchet surface 218' of the first ratchet portion 214 is urged against the rearward facing ratchet surface 218" of the second ratchet portion 216. Accordingly, when driven, the thrust of the fan 208 automatically acts to urge the ratchet mechanism 206 together, thereby maintaining adequate power transmission through the ratchet mechanism from the second shaft portion 204 to the first shaft portion 202 (and thus the fan 208).
  • the air resistance on the fan 208 urges it and the first shaft portion 202 rearward.
  • the force on the fan 208 urges the first ratchet portion 214 away from the second ratchet portion 216. Accordingly, when windmilling, the air resistance on the fan 208 automatically acts to urge the ratchet mechanism 206 apart, thereby ensuring that the first shaft portion 202 and the fan 208 can rotate freely with respect to the second shaft portion 204.
  • the ratchet mechanism 206 is configured such that downstream axial movement of or axial force on the first shaft portion 202 relative to the second shaft portion 204 causes the ratchet surfaces 218',218" of the first and second ratchet elements 214,216 to be disengaged.
  • the shaft apparatus 200 described herein utilises the inherent forces in the system in the driven and windmill conditions to optimise the operation of the ratchet mechanism 206.
  • FIGS 8a and 8b show an exemplary ratchet surface configuration suitable for use in the shaft apparatuses described herein.
  • the ratchet surfaces 318' and 318" each comprise a plurality of complementarily shaped ratchet teeth 320.
  • the ratchet teeth 320 are configured such that they interlock in one direction only.
  • the ratchet surface 318' is moving in the direction illustrated by arrow A. Owing to the interlocking of the ratchet teeth 320 of surface 318' with the ratchet teeth 320 of the other ratchet surface 318", the second ratchet surface 318" is driven in the same direction A.
  • This is comparable to a driven condition of the fan, in which the ratchet surface 318' is of the second shaft portion is applying a driving torque to the ratchet surface 318", which of the first shaft portion.
  • the ratchet teeth have an undercut such that there is no separation of the teeth when torque loaded.
  • Figure 8b shows an alternative condition, in which the ratchet surface 318' is static, and the ratchet surface 318" is moving in the direction A.
  • condition in this illustration could also represent a condition in which both surfaces 318' and 318" are moving in direction A, but surface 318" is moving faster than surface 318'.
  • the ratchet teeth 320 do not interlock such that the ratchet teeth 310 of the second ratchet surface 318" can freely ride past the ratchet teeth 320 of the first ratchet surface 318'.
  • This is comparable to a windmill condition of the fan, in which the ratchet surface 318' is of the second shaft portion which has seized, and the ratchet surface 318" is of the first shaft portion which is rotating with the windmilling fan.
  • the ratchet teeth 320 have a friction surface 322 which is configured to maintain frictional engagement in a simple 'power-loss' windmill condition when the second shaft portion is still able to rotate, but slide over one another to allow free rotation of the first shaft portion when the second shaft portion is seized.
  • FIG 9 shows a curvic coupling 400 suitable for use as the first and second ratchet elements of the shaft apparatus.
  • the curvic coupling 400 presents an annular ratchet surface 402 for engagement with a corresponding curvic coupling.
  • the ratchet surface 402 has a plurality of circumferentially spaced curvic teeth 404.
  • Each of the curvic teeth 404 is radially curved. Accordingly, the curvic teeth 404 act inherently to centre the curvic coupling 400 with a corresponding curvic coupling, which ensures that the shaft portions to which the couplings 400 are attached maintain alignment under load despite not being fixedly connected.
  • shaft apparatuses described herein permit a method of operating a gas turbine engine in which a fan is permitted to rotate relative to a torque shaft in a first direction, and the fan is prevented from rotating relative to the torque shaft in a second direction opposite to the first direction.
  • an engine fan can be driven normally by a torque shaft.
  • the fan is permitted by the ratchet mechanism to windmill relative to the torque shaft.
  • the drag of the engine can be minimised in an engine failure scenario by virtue of the shaft mechanism, maintaining better aircraft control.
  • the principles of the present disclosure may be particularly advantageous in a gas turbine engine having a power gear box, as gearboxes might provide additional potential failure routes which could result in a torque shaft seizure.
  • Some of the embodiments described herein provide additional advantages in exploiting the reversal of axial forces in the engine system between driven and windmill conditions to automatically decouple the first and second shaft portions when required.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • General Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP20157725.1A 2019-03-06 2020-02-17 Wellenvorrichtung für ein gasturbinentriebwerk Active EP3705704B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GBGB1902980.0A GB201902980D0 (en) 2019-03-06 2019-03-06 Shaft apparatus for a gas turbine engine

Publications (2)

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EP3705704A1 true EP3705704A1 (de) 2020-09-09
EP3705704B1 EP3705704B1 (de) 2022-04-20

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Citations (4)

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Publication number Priority date Publication date Assignee Title
GB866046A (en) * 1958-04-09 1961-04-26 United Aircraft Corp Clutch mechanism for connecting a rotary driving member with a rotary driven member
US20160003143A1 (en) * 2013-02-26 2016-01-07 United Technologies Corporation Turbomachine fan clutch
EP3002476A1 (de) * 2014-10-02 2016-04-06 Siemens Aktiengesellschaft Kupplung von zwei Wellen mit mechanisch vorgegebenem Kuppelwinkel und zugehöriges Kuppelverfahren
EP3396115A1 (de) * 2017-04-28 2018-10-31 Rolls-Royce Deutschland Ltd & Co KG Gasturbinentriebwerk

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Publication number Priority date Publication date Assignee Title
GB744923A (en) 1953-05-19 1956-02-15 Rolls Royce Improvements relating to gas turbine engines
GB790550A (en) 1955-01-25 1958-02-12 Rolls Royce Improvements relating to aircraft gas-turbine engines
US5205386A (en) 1992-03-05 1993-04-27 United Technologies Corporation Pawl and ratchet clutch with pawl holdback
US7334392B2 (en) * 2004-10-29 2008-02-26 General Electric Company Counter-rotating gas turbine engine and method of assembling same
CN103486154B (zh) 2013-09-30 2015-12-23 中国科学院工程热物理研究所 燃气涡轮发动机起动用超越离合器
WO2019092910A1 (ja) 2017-11-13 2019-05-16 株式会社Ihi ターボファンエンジン

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB866046A (en) * 1958-04-09 1961-04-26 United Aircraft Corp Clutch mechanism for connecting a rotary driving member with a rotary driven member
US20160003143A1 (en) * 2013-02-26 2016-01-07 United Technologies Corporation Turbomachine fan clutch
EP3002476A1 (de) * 2014-10-02 2016-04-06 Siemens Aktiengesellschaft Kupplung von zwei Wellen mit mechanisch vorgegebenem Kuppelwinkel und zugehöriges Kuppelverfahren
EP3396115A1 (de) * 2017-04-28 2018-10-31 Rolls-Royce Deutschland Ltd & Co KG Gasturbinentriebwerk

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GB201902980D0 (en) 2019-04-17
EP3705704B1 (de) 2022-04-20
US20200284202A1 (en) 2020-09-10

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